This invention relates generally to hybrid rockets, and, more particularly to a constant thrust hybrid rocket motor capable of maintaining operation over a wide temperature range.
An excellent example of a hybrid rocket motor is set forth in U.S. Pat. No. 3,156,092 by this inventor. As described therein the hybrid rocket motor is a small, easily transportable rocket motor which utilizes a safe combustion system operating with safe propellants. It is completely restartable due to its self contained ignition system and the motor combustion can be viewed through the actual propellant fuel. Therefore, it is especially useful as a laboratory tool for the evaluation of chamber configuration, oxidizer flow rate, grain length, grain composition and the like.
More specifically, the hybrid rocket motor includes a combustion chamber of an oxidizable semi-transparent plastic such as Plexiglas (polymethylmethacrylate, also known as Lucite), polystyrene, polyethylene, Teflon, polybutyrate or the like. The plastic serves both as the actual fuel during operation and also as a transparent chamber so that one can watch the action of the hybrid rocket motor. Plexiglas is ideal for this purpose since it serves as a typical hybrid fuel when used with oxygen and is clean-burning, and does not emit a carbonaceous exhaust. Furthermore, it is a thermal plastic which is readily available as a molding powder as well as in rods and tubes, so that combustion chambers of any desired configuration can be fabricated without difficulty. Since oxygen and plastic are not hypergolic, an electric ignition system can form part of the hybrid unit.
Unfortunately, the hybrid rocket motor is subject to a wide temperature range of, for example, -65.degree. F. through +145.degree. F. As a result of this temperature difference the pressure variation in the oxygen tank can range from between 2500 to 5500 psia. In fact, after passing through a constant pressure regulator, the temperature extremes drop to -200.degree. F. and 22.degree. F. respectively at 200 psi. Even the utilization of a feedback signal from a transducer to open and close a main solenoid valve feeding oxidizer into the combustion chamber still results in variations in chamber pressure of .+-.10-15%.
It is therfore clearly evident from the above description of the drawbacks associated with prior art hybrid rocket motors, that it would be highly desirable to provide a constant thrust hybrid rocket motor which would be capable of operating over wide temperature ranges of, for example, -65.degree. F. through +145.degree. F. without experiencing the pressure variations normally associated with such prior hybrid rocket motors.